Minimize FormoSat-5


The FormoSat-5 (FS-5) program represents the first indigenous development of remote sensing capabilities at NSPO (National SPace Organization) of Taiwan. After successfully conducting three satellite missions during in the timeframe 1991-2006, NSPO started a new space initiative in 2004 which emphasizes on building up the capabilities for independent development of spacecraft and payload instruments. The strategy for the program is to acquire key technology and setup a heritage bus design for NSPO. The key technology to be developed by NSPO includes flight software, EGSE and IPS. 1) 2) 3) 4) 5) 6)

FormoSat-5 is a follow-on mission of FormoSat-2. An important objective for FormoSat-5 is to build up the domestic capability for the high-resolution optical remote sensing instrument by integrating NSPO with the ITRC (Instrument Technology Research Center) of NARL (National Applied Research Laboratories). Both NSPO and ITRC belong to the NARL, a non-profit organization under the supervision of NSC (National Science Council).

Development strategy: In the past two decades, NSPO has successfully finished three space programs mainly through prime contractors from either USA or Europe. NSPO engineers have been broadly involved in system & subsystem designs and became familiar with the systems engineering practices utilized by both European and USA aerospace companies. - For the FormoSat-5 program, NSPO has decided to take a bold step on taking up the full responsibility on satellite systems/subsystems analyses/designs, component development, and payload/satellite integration and test, as shown in Figure 1. 7)

This program is the first self-reliant space program in Taiwan and is serving as a launching platform to domestically build important spacecraft components. Under the guidance of NARL, NSPO has integrated high technology industries, research institutes and companies into a competent team to execute the program. Key team players include CSI (COMOS Sensor Inc.), Camels, ITRC (Instrument Technology Research Center), the National Chip Implementation Center (CIC), AIDC (Aerospace Industrial Development Corporation), and CSIST (ChungShan Institute of Science and Technology).

In addition, NSPO obtains outsourcing components through STI (SpaceTech GmbH, Immenstaad), Germany. SpaceX (Space Exploration Technologies Corporation) will provide the launch service.


Figure 1: Overview of the FormoSat-5 development strategy (image credit: NSPO)



NSPO started with the development of the FormoSat-5 project in 2005. The design strategy of the spacecraft bus includes the following major considerations: 8) 9)

1) To develop an NSPO heritage platform architecture capable of extending to the future high-resolution RSS (Remote Sensing Satellite).

2) To develop the avionics, flight software, and control algorithms extendable and reusable to the future missions

3) To achieve cost effective designs through revising product assurance strategies and acquiring COTS (Commercial Off The Shelf) components.

The spacecraft bus is shown in Figure 2. FormoSat-5 is a minisatellite of octagonal shape (3 m in height and 1.2 m in diameter) with a wet mass of ~ 525 kg. The spacecraft design life is 5 years. The spacecraft is 3-axis stabilized. The design provides for an agile body-pointing capability of the spacecraft with a FOR of ±45º (cross-track or along-track).


Figure 2: Accommodation of the spacecraft bus components (image credit: NSPO)


Figure 3: Illustration of the deployed FormoSat-5 spacecraft configuration (image credit: NSPO)


Figure 4: Configuration of the system architecture (image credit: NSPO)

The CDMU (Command and Data Management Unit) is used for spacecraft control. It is composed of the following functional elements:

- OBC (Onboard Computer) is a Leon3-FT (Fault Tolerant ) processor implemented on an Actel AX-2000 FPGA. The processor is highly configurable, and includes functions such as the Floating Point Unit, three SpaceWire links, and a memory controller. The board also features UART (Universal Asynchronous Receiver/Transmitter) ports, CAN (Controller Area Network) bus controller and a clock system. The OBC itself also features the capability of board-level redundancy control of the other unit.

The LEON-3 CPU, developed by Gaisler Research, is a 32 bit synthesisable processor core based on the SPARC V8 architecture. It was chosen because of its advantages such as its high level of configurability and fault immunity, the wide availability of the source code, a deep pipeline and multi-processor support, and its suitability for the SOC (System-On-Chip) design through its implementation in a flight qualified FPGA (Field Programmable Gate Array).

- DS (Data Storage) module. Used to store satellite health telemetry and scientific data. DS has a memory capacity of 256 MB SDRAM with EDAC protection. Among other interfaces, it comprises also two SpaceWire nodes with CRC protection.


Figure 5: Overview of the CDMU architecture (image credit: NSPO)

- Reconfiguration Electronics (RE) module. The objective is to manage the redundancy of the CDMU. The module monitors the major alarms of the CDMU; in case of a system irregularity, the module is able to reset the current CDMU, or switch over to the other CDMU, or employ a cross-strapping configuration.

- TMTC (Telemetry and Telecommand) module. The TMTC module encodes/decodes the telemetry/telecommand data according to the ESA packet telemetry standard. There are two

- Housekeeping (HK1, HK2) module. The objectives are to acquire the housekeeping data and to control satellite components. There are two identical modules in each CDMU. Each housekeeping module provides analog inputs and outputs, digital inputs and outputs, and serial interfaces. The BC communicates with two GPSR modules via housekeeping modules.

- AOCS (Attitude Orbit and Control Subsystem) module. The implementation includes the following interfaces: 2 star cameras, 2 magnetometers, 6 sun sensors, 4 reaction wheels, 3 magnetic torquers, and 2 pressure transducers.

- SpaceWire Mux-Demux (Multiplexer-Demultiplexer), developed by Carlo Gavazzi Space S.p.A., Italy. It features two master nodes connected to OBC-A and OBC-B, and 6 slave nodes connected to HK1, HK2, AOCS, DS, TMTC and one spare node. The data rate used for CDMU is 50Mbit/s.

- GPS receiver (GPSR) modules. Two GPS receiver modules are used in the CDMU. One GPS receiver is selected as the master and the second one as the backup receiver. Each GPSR module has its own PPS signal, however just one is selected. The PPS (pulse per second) signal shall be used for synchronization of the OBC clock which is distributed within the satellite.

- DC to DC power supply (DCDC1 & DCDC2) modules. The DCDC1 and DCDC2 convert the external unregulated 28V power to stable secondary powers for the CDMU modules. The DCDC1 module provides power to TMTC, SpaceWire MUX-DEMUX, RE and backplane modules. The DCDC2 module provides power to BC, DS, GPSR1, GPSR2, HK1, HK2, and AOCS modules.


Figure 6: SpaceWire Mux-Demux architecture (image credit: NSPO)

EPS (Electrical Power Subsystem): The EPS is composed of a PCDU (Power Control Distribution Unit), a rechargeable battery, and two deployable solar generators. The PCDU is designed to handle over 150 FET switch outlet channels, converts the primary 28 V to the users demand secondary voltage, as well as regulates solar power input to the spacecraft. The FS-5 battery has a capacity of 24 Ah provided by small cells in a 8 series 16 parallel configuration. Each solar panel contains of 19 solar strings with 20 TJ GaAs solar cells in series. The solar generators can support FS5 mission in 280 W orbit average power.

The PCDU was domestically developed. The device contains seventeen circuitry modules with more than a hundred power distribution channels. Power is collected from 38 solar electricity control paths. The PCDU EM has been intensively tested on the satellite EDM (Engineering Development Module) along with the OBC EM and an early-released flight software build to verify the satellite command and telemetry contents. 10) 11)

The main functions of the PCDU are to condition energy from the solar arrays and distribute power for all subsystems on the satellite. The PR (Power Relay) module, one of the major modules in the PCDU, is to manage the switchover of redundant DC-DC converter sections, secondary power distribution and power protections. Each set of internal/external DC-DC converter section which is protected by an OC (Over Current) protection with LCL (Latch-up Current Limiter) is controlled by the PR module’s FPGA. To avoid a short circuit in the DC-DC converter sections, the PR module provides UV (Under Voltage) protection to switch the nominal DC-DC converter to the redundant DC-DC section. In order to monitor the primary power voltage, an UVLO (Undervoltage-Lockout) detection circuit with majority voting in the FPGA is integrated in PR module. Figure 7 shows the main functional blocks of the PR module.


Figure 7: Functional block diagram of the PR module (image credit: NSPO)


Figure 8: Block diagram of the PCDU (image credit: NSPO)


Figure 9: Component layout of the FormoSat-5 spacecraft (image credit: NSPO)

RF communications: The X-band is used for payload data downlink with a data rate of up to 150 Mbit/s. The TT&C functions are provided in S-band. An onboard image storage capability of 80 Gbit is provided. 12)

Key parameter



Orbit, SSO (Sun Synchronous Orbit)

720 km

891 km

Revisit time (capability)

Every other day


LTDN (Local Time of Descending Node)

9:45-10:15 hours

9:30-10:30 hours

Spacecraft design life

5 years

5 years


> 0.6 as a design goal at EOL

> 0.6 at EOL

GSD (Ground Sample Distance)

2 m (Pan), 4 m (MS)

2 m (Pan), 8 m (MS)

Swath width

24 km

24 km

Spectral bands

1 Pan + 4 MS

1 Pan + 4 MS

System CTF (Contrast Transfer Function), Pan

≥ 0.1

≥ 0.12

SNR (Signal-to-Noise Ratio), Pan

≥ 92

≥ 92

Duty cycle of RSI (Remote Sensing Imager)

≥ 8% per orbit

≥ 8% per orbit

FOR (Field of Regard)



Spacecraft agility

Roll : 24º/60 s
Pitch : 24º/60 s
Yaw : 7º/60 s

Roll : 60º/60 s
Pitch : 50º/60 s
Yaw : 35º/60 s

Pointing accuracy

0.1º per axis

0.1º per axis

Geo-accuracy of imagery

With GCP: 2 m
W/O GCP: 390 m

With GCP: 2 m
W/O GCP: 470 m

On-board storage capacity

80 Gbit

45 Gbit

Downlink availability for image data

≥ 7 orbits/day

≥ 7 orbits/day

Spacecraft mass

~525 kg

764 kg

Table 1: Comparison of key parameters for FormoSat-5 and FormoSat-2


Launch: A launch of FormoSat-5 is scheduled for 2015 on a Falcon-1e vehicle of SpaceX, Hawthorne, CA, USA. The launch site is Omelek Island at the U.S. Army Kwajalein Atoll (USAKA) in the Central Pacific (about 3900 km southwest of Honolulu, the Kwajalein Atoll is part of the Republic of the Marshall Islands). A launch contract between NSPO and SpaceX was signed in June 2010. 13) 14)

Orbit: Sun-synchronous near-circular orbit, altitude = 720 km, inclination = 98.28º, period = 99.19 minutes, LTDN (Local Time on Descending Node) at ~ 10 hours.



Sensor complement: (RSI, AIP)

RSI (Remote Sensing Imager):

RSI, the primary payload, is being developed domestically by a team including NSPO institutions and commercial partners: ITRC (Instrument Technology Research Center), CIC (National Chip Implementation Center), CSIST (Chung-Shan Institute of Science and Technology), AIDC (Aerospace Industrial Development Corporation), CIS (CMOS Sensor Inc.), and CAMELS Vision Technologies.

NSPO/ITRC/CIC of NARL (National Applied Research Laboratories) coordinate the domestic multi-discipline capabilities to collaborate for developing the fully made-in-Taiwan (MIT) RSI payload and establishing the background for the future higher performance RSI instrumentation. 15) 16) 17) 18)

The objective is to establish capabilities of system design, integration & test, and opto-mechanical system calibration. The goal is to develop a self-reliant infrastructure in all key technologies of the project. The imaging objectives are to provide a 2 m GSD (Ground Sample Distance) for Pan imagery and 4 m GSD for multispectral imagery on a swath of 24 km.

The imagery of RSI will be used for such applications as: planning tool for urban development, environment and crop health monitoring, disaster assessment, etc.


Figure 10: Illustration of the RSI (image credit: NSPO)

RSI instrument type

Pushbroom imager

Spectral bands

1 Pan
4 MS (Multispectral) bands: Blue, Green, Red, NIR

GSD (Ground Sample Distance)

2 m (Pan), 4 m (MS)

Swath width

24 km

Data quantization

12 bit

CTF (Contrast Transfer Function)

≥ 0.1

SNR (Signal-to-Noise Ratio)

≥ 92 (Pan), ≥ 100 (MS)

Detector type

TDI CMOS array, 5 bands, Pan =12,000 pixels, MS = 6000 pixels


Cassegrain telescope with an aperture of 45 cm (primary mirror)

Table 2: Parameters of the RSI instrument

The RSI instrument consists of a Cassegrain telescope and an EU (Electronic Unit). The telescope is composed of the primary mirror (entrance pupil diameter = 45cm), secondary mirror, correct lens, baffles, CFRP (Carbon Fiber Reinforced Plastic) structure, and FPA (Focal Plan Assembly ). The FPA is the most crucial component in terms of its technology and performance for space applications (Ref. 5).


Figure 11: Schematic view of the RSI telescope (image credit: NSPO)

The FPA is composed of the image sensor, bandpass filter, control electronics, and the housing structure. Detector arrays with CMOS (Complementary Metal Oxide Semiconductor) technology are used. The Pan detector is composed of 12,000 pixels with 10 µm/pixel and the MS image sensor is composed of 6,000 pixels with 20 µm per pixel, respectively, that provide the spatial resolutions of 2 m (Pan) and 4 m (MS) in a swath of 24 km. The detector array chip of CIS (CMOS Sensor Inc.) has a size of 12 cm x 2.4 cm. This advanced detector array has been designed by CIS and fabricated on an 8” wafer with 0.18 µm process manufactured by UMC (United Microelectronics Corporation) in Taiwan.


Figure 12: Photo of a test chip of CIS (left) and the detector array on an 8” wafer (right), image credit: NSPO

A 5-band filter has been designed and developed in-house, meeting (actually exceeding) all performance requirements for RSI integration (Figure 13).


Figure 13: Illustration of the filter and its spectral band performance (image credit: NSPO, Ref. 7)

The acquisition of on-board image data compression techniques represents another achievement for this program. Collaboration between NSPO and its vendor has mastered the CCSDS format through applying DWT (Discrete Wavelet Transfer) and BPE (Bit Plane Encoder) techniques to compress the very high volume imagery in real time.


Figure 14: Overview of the RSI system architecture (image credit: NSPO)

On June 11, 2012, the FormoSat-5 program held the AIT (Assembly, Integration, and Test) kick off meeting at NSPO for the RSI (Remote Sensing Instrument). 19)


AIP (Advanced Ionospheric Probe):

The AIP instrument is designed and developed at NCU (National Central University), Jhongli City, Taiwan. It is an all-in-one plasma sensor with a sampling rate up to 8,192 Hz to measure ionospheric plasma concentrations, velocities, temperatures, and ambient magnetic fields over a wide range of spatial scales. With the comprehensive dataset from the AIP observations, the project will be able to conduct a systematic examination of the plasma parameters in the topside F region of the ionosphere, with regard to longitudinal and latitudinal distributions and seasonal variations. The collected data will not only be used for space science studies, but can also be shared by the scientific community in seismic precursors associated with strong earthquakes and co-seismic effects.

The AIP instrument is being developed through a collaborative effort between scientists of ISS (Institute of Space Science) and engineers of IOM (Institute of Opto-Mechatronics) at NCU, led by Chien-Ming Huang. These two teams have participated in many space payload missions, funded by NSPO, and in international collaborations, providing the scientific payloads. The payloads obtained their flight heritage on sounding rocket flight and on spacecraft throughout the last decade.

Note: In the early stage of the program planning, the two instruments MFI (Magnetic Field Instrument) and MDRP (Marine Data Relay Payload) were among the scientific payload candidates for the FormoSat-5 mission. However, since the later part of 2011, AIP (Advanced Ionospheric Probe) has become the only scientific payload on board FormoSat-5. 20)

The AIP instrument features measurement modes like a PLP (Planar Langmuir Probe), RPA (Retarding Potential Analyzer), an IT (Ion Trap), and an IDM (Ion Drift Meter) to obtain the characteristics of the ionospheric plasma concentration (Ni), velocities (Vi), and temperatures (Ti and Te). 21)

1) PLP: When AIP operates in PLP mode, the PLP electrode (shown on the left of Figure 15) is isolated with aperture plane and applied to a sweeping voltage between -10 V and +10 V. Once the sweeping voltage is lower than the plasma potential, the electrode repels electrons and accelerates ions. As the voltage decreases, the electron current decreases rapidly. The electron temperature (Te) can be determined by the slope of the I-V curve by the following equation:


where I is the current through the PLP electrode, V the voltage applied to PLP electrode, κ is the Boltzmann constant, and e is the electric charge. As the slope of the I-V curve is flat, the Te is higher. As the slope is steep, the Te is lower.


Figure 15: Schematic view of the AIP operation in the RPA and IT mode (image credit: NCU)

2) RPA and IT: AIP consists of three metal grids to handle the incoming plasma (Figure 15). The first grid, G1, is an aperture grid arranged in the front of the probe and connected to a floating potential device to allow a smooth entry of the incoming plasma into the probe, not to be affected by the transverse electric field. The second grid, G2, is a retarding grid connected to a programmable sweeping voltage circuit between -10 and 10 V. If AIP operates in the IT mode, the grid is maintained at the floating potential without repelling the incoming ions. In the RPA mode, the grid forms a potential barrier (retarding voltage) to block the low energy positive ions into the probe. The third grid, G3, is a suppressor grid, maintained at -15 V to repel the incoming electrons into a quadrant metal collector; it can also reduce the photoelectrons escaped from the collector when the sunrays strike on the collector. All four metal plates of the quadrant collector are maintained at the floating potential and are connected together to an ammeter via an analog switch in these two modes.

As the retarding voltage increases, the positive ions with low kinetic energy are repelled out of the probe or neutralized by the grids and interior boundary of the probe. The positive ions with high kinetic energy can penetrate through the retarding grid to the collector if they do not contact with the suppressor grid or the interior boundary of the probe. Basically the higher the retarding voltage, the lower the current is measured at the collector. Such a sampling process can be recorded as I-V curves that are embedded with information of ion temperature, composition and ram velocity. A 1-dimensional ion current equation for this application has been written down in the following form:


where I is the current measured at the collector, q is the electric charge, A is the effective collection area of the collector, n is the ion concentration, U is the ion ram velocity,M is the mass of the ion, κ is the Boltzmann constant, Ti is the ion temperature, Vr is the maximum electric potential along the path for plasma from outside environment to the probe (it is usually the retarding voltage). It should be noted that the electric potential is referenced to far-away plasma (the electric potential of the far-away plasma is zero).

The initial ion temperature and the concentrations can be deduced from the I-V curves via a half-current approximation. To improve the quality of the output parameters, an iterated half-current method is prepared for initial guesses and a grid search method is used to refine the output parameters with possible uncertainties. To derive precise ion temperature and ram velocity, a numerical model to estimate the effects of the grid alignment and electric potential depression on the grids is required.

3) IDM: In IDM mode, the G1 and G2 grids are maintained to the floating potential. However, each metal plate of the quadrant collector is connected to an individual current meter (ammeter). The arrival angle of the incoming ions can be estimated from the current differences of the adjacent ammeters (Figure 16). The possible current ratio can be estimated by the following relation:


where W is the width of the aperture, D is the depth between the aperture and the collector. The incident angle, , of the incoming ion can be expressed as:


Figure 16: Illustration of the AIP operation in the IDM mode (image credit: NCU)

4) Geophysical Parameters: In default mode, the AIP is set to measure the ionospheric plasma in a cycle of PLP, RPA, IT, and IDM mode sequentially. It is assumed that all ionospheric parameters have a slow variation within a 4-second period. After one cycle is complete, the geophysical parameters, Ni, Vi, Ti, and Te, can be derived from the steps shown in Figure 17.


Figure 17: Scheme of the AIP data flow to derive the geophysical parameters (image credit: NCU)

Vi can be obtained by the arrival angles,θH and θV, from the IDM mode and U from the RPA mode. Ni can be obtained by ion flux from IT mode and U from the RPA mode. Ti and U can be obtained by the I-V curve from the RPA mode, and Te from the PLP mode.


AIP instrument: The AIP consists of two main units, a sensor unit located on the top panel of the satellite and a SPEU (Science Payload Electronics Unit) inside the satellite. The sensor unit has a sensor head mounted on a stand to enlarge the FOV (Field of View). The SPEU hosts the controllers and the DC/DC converters to control the sensor unit and interface with the CDMU (Command and Data Management Unit) of the satellite bus.


100 mm (L) x 100 mm (W) x 100 mm (H) for the sensor head
140 mm (L) x 130 mm (W) x 380 mm (H) for the sensor stand
180 mm (L) x 180 mm (W) x 60 mm (H) for the SPEU


0.582 kg for the sensor head
0.916 kg for the sensor stand
~ 2 kg for the SPEU
~ 1 kg for the harness


Al alloy 6061-T6 for the sensor head
Al alloy 7075-T6 for the sensor stand
Al alloy 6061-T6 for the SPEU

Table 3: Mechanical specification of the AIP

The sensor head is illustrated in the Figure 18 and its three-grid configuration is shown in Figure 19. The FOV, with the angle centered velocity vector (sensor look direction) and originating at the perimeter edge of the AIP aperture plane, is about 56º. The opening of the sensor is 50 mm. All the grids are made of stainless steel 316, coated with gold, with 50 lines per inch in grid density and 0.0045 inch in wire diameter.


Figure 18: The 3-D model of the AIP sensor head (image credit: NCU)


Figure 19: The profile of the AIP sensor head (image credit: NCU)

Electrical specification: The functional diagram of the AIP is shown in the Figure 20. Dual CDMUs and a PCDU (Power Control Distribution Unit) with dual power channels are interfaced with the SPEU. A redundant design is provided for the controllers, DC/DC converters, A/D converters, and D/A converters; the ammeters are implemented to reduce a single point failure of the system. Since there are no hardwires from the PCDU to activate the primary or redundant controller, the PCDU can provide power to the primary/redundant DC/DC converter to startup the primary/redundant controller. However, there is only one controller on at a time. It is forbidden to provide power to the primary and the redundant DC/DC converters simultaneously. This would cause a cross-strapping malfunction of the RS-422 interfaces to the CDMU and digital lines to the sensor.


Figure 20: Functional block diagram of the AIP (image credit: NCU)

Input voltage

28±6 VDC

Average power

2.24 W for sensor, ~2.5 for SPEU

Command and telemetry interface

Dual channels, RS-422 serial asynchronous transmission with a baud rate of 19.2 kbit/s

Science data interface

Dual channels, RS-422 serial synchronous transmission with a baud rate of 1 Mbit/s

Table 4: Electrical parameters of the AIP

In the command and telemetry interface, 3-byte commands and 64-byte ancillary data are designed to deliver from the CDMU to the SPEU. The 19-byte status of health information of the AIP is delivered to from the SPEU to the CDMU.

In the science data interface, a science data packet with a fixed length of 1024 bytes can be delivered from the SPEU to the CDMU. For the normal/fast/burst mode, the SPEU can deliver 1/8/64 science data packet(s) every 3 seconds.

Ion composition (light ions and O+)

Range: 3% to 100%
Sensitivity: < 1%
Accuracy: < 10%
Duration for a sample: 1 s

Total ion concentration

Range: 4 x 102 to 1.2 x 107 # cm-3
Sensitivity: < 1%
Accuracy: < 10%
Duration for a sample: 1 s

Ion velocity

Range: ±4 km/s for cross track direction, ±5 km/s for ram direction
Sensitivity: ±10/s for cross track direction, ±100 m/s for ram direction
Accuracy: ±50 m/s for cross track direction, ±200 m/s for ram direction
Duration for a sample: 1 s

Ion temperature

Range: 500 to 10,000 K
Sensitivity: ±50 K
Accuracy: ±200 K
Duration for a sample: 1 s

Electron temperature

Range: 500 to 10,000 K
Sensitivity: ±50 K
Accuracy: ±200 K
Duration for a sample: 1 s

Table 5: Parameter limits of the AIP data



Ground segment:

Ground segment, which includes ground operation facility and image processing system, is being upgraded to meeting FormoSat-5 requirements. Both hardware and software upgrades are being conducted on top of the existing operational segment by NSPO team in house with minimum external supports.


Figure 21: Overview of the FormoSat-5 system (image credit: NSPO)

1) H. P. Chang, “FormoSat-5 Program,” NSPO, Dec. 9, 2010, URL:

2) Albert Lin, Chien-Fang Lai, “The Command and Data Management Unit for NSPO FormoSat-5 and Future Satellites,” 4th Asian Space Conference (ASC-2008), Taipei, Taiwan, October 1-3, 2008, URL:

3) Jeng-Shing Chern, Jer Ling, Shui-Lin Weng, “Taiwan's second remote sensing satellite,” Acta Astronautica, Vol. 63, Issues 11-12, December 2008, pp. 1305-1311

4) Bang-Ji Wang, Chia-Ray Chen, Jih-Run Tsai, Fei-Hon Chou, “Remote Sensing Instrument Development in Taiwan,” 4th Asian Space Conference (ASC-2008), Taipei, Taiwan, October 1-3, 2008

5) Guey-Shin Chang, An-Ming Wu, Ho-Pen Chang, “A Perspective on Taiwan’s Earth Observation Missions,” Proceedings of the 63rd IAC (International Astronautical Congress), Naples, Italy, Oct. 1-5, 2012, paper: IAC-12-B1.2.6

6) Lou-Chuang Lee, “Space Programs in Taiwan - FORMOSAT 5 and FORMOSAT 7,” 4th International Conference on Particle and Fundamental Physics in Space (SpacePart12),” CERN, Geneva, Switzerland, Nov. 5-7, 2012, URL:

7) Ho-Pen Chang, Guey-Shin Chang, Jer Ling, Tony Tsai, “Remote Sensing Satellite FORMOSAT-5,” Proceedings of the 63rd IAC (International Astronautical Congress), Naples, Italy, Oct. 1-5, 2012, paper: IAC-12-B4.4.8

8) Jeng-Shing Chern, Jer Ling, Shui-Lin Weng, “Taiwan’s second remote sensing satellite,” Acta Astronautica, Vol. 63, Issues 11-12, December 2008, pp. 1305-1311

9) “FORMOSAT-5 Component Environmental Specification,” NSPO, FS5-SPEC-0003, Issue 01 Revision 00, April 21, 2010, URL:

10) Che-Cheng Huang, Jia-Jing Yeh, Zhe-Yang Huang, Chien-Kai Tseng, “FormoSat-5 Satellite Power Protection Design,” Applied Mechanics and Materials,Vol. 145, 2012, pp. 536-541


12) I-Young Tarn, “On the Antenna Gain Requirement Derivation and Antenna Orientation Arrangement for Satellite Remote-Sensing Data Downlink,” Proceedings of the 4th Asian Space Conference (ASC2008) and the FORMOSAT-3/COSMIC Data User Workshop, Taipei, Taiwan, Oct. 1-3, 2008

13) “SpaceX And NSPO Sign Contract To Launch Earth Observation Satellite,” Space Daily, June 17, 2010, URL:

14) Peter B. de Selding, “SpaceX Falcon 1e To Launch Taiwan’s FormoSat-5 Craft,” Space News, June 15, 2010, URL:

15) Jer Ling, “Applications of CMOS Sensors on FORMOSAT-5 Remote Sensing Instrument,” Workshop on CMOS Applications in Astronomy and Space Sciences, National Tsing Hua University, Hsinchu, Taiwan, Jan. 5-6, 2011, URL:

16) “Remote Sensing Instrument Made Domestically - NSPO Spreads the Wings to Fly Again,” NSPO, Feb. 2, 2010, URL:


18) Wen-Chih Hsu, Jer Ling, Wei-Chun Chen, Ming-Hsien Tsai, Shih-Hung, “The Radiation Qualification of the Taiwanese CMOS Image Sensor for the Remote Sensing Satellite,” 8th International Hiroshima Symposium of the Development and Application of Semiconductor Tracking Detectors. Taipei, Taiwan, Dec. 8, 2011, URL:

19) “FORMOSAT-5 kicked off Remote Sensing Instrument AIT Tasks,” NSPO, June 12, 2012, URL:

20) Information provided by Ho-Pen Chang, FormoSat-5 Program Director of NSPO.

21) Chi-Kuang Chao, Chien-Ming Huang, Yen-Hsyang Chu, Ching-Lun Su, Jann-Yenq Liu, Shigeyuki Minami, “Advanced Ionospheric Probe onboard the FORMOSAT-5 Satellite,” Proceedings of the UN/Japan Workshop and The 4th Nanosatellite Symposium, Nagoya, Japan, Oct. 10-13, 2012

The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.